Spacecraft and System Description

The Northrop-Grumman Shepherd Spacecraft (S-S/C) is designed to direct the impact of the LRO Centaur upper stage within 10km of a target inside a permanently shadowed region at the North or South Pole of the Moon with high reliability. The S-S/C configuration reduces development and performance risk to its own mission, as well as the LRO mission, by using existing, flight-proven avionics, propulsion elements, structures, adapters, and release systems, all in a high-frequency structural configuration.

After launch, the Centaur performs its trans-lunar-injection burn and then releases LRO. Following this, the Centaur performs an LRO divert maneuver and is targeted to provide a 70 m/s delta-V that positions the Centaur & S-S/C stack for a lunar-gravity-assisted fly-by of the Moon, providing a separation trajectory between LRO and the Centaur & S-S/C stack. Following this burn and two hours into the mission, the S-S/C takes trajectory responsibility for the Centaur & S-S/C stack while LRO proceeds to the Moon. The Centaur conducts a propellant-vent sequence, leaving its primary engine valve open while it remains passive for the remainder of the mission. This venting procedure reduces the left-over propellent mass in the Centaur fuel tanks. The hydrogen and oxygen boil-off over 84 days eliminates the possibility of either being present in the impact plume at the mission end. After a lunar flyby and 86 days after launch, the S-S/C performs a final targeting maneuver to line up the Centaur for impact. The S-S/C separates from the deactivated Centaur it shepherded towards impact, and performs a delay burn that causes it to follow the Centaur by 10 minutes. The S-S/C inverts and then operates as an in-situ observer of the plume when Centaur impacts. Fifteen minutes later, the S-S/C also impacts creating a smaller, Earth-observable plume. The remainder of this section defines the Shepherd Spacecraft and its mission operation in detail.

Shepherd Spacecraft Interface with Launch Vehicle and LRO
The LCROSS mission has minimal physical and functional impact on the LRO mission. The LCROSS S-S/C and its adapter are inserted between the Centaur and the LRO adapter. LCROSS S-S/C’s 62” interface replicates the Centaur interface for the LRO-Centaur adapter. The Northrop-Grumman team will implement a systems engineering approach to minimize any issues or risks that could derive from this new structure including; analytical (e.g., finite element models), mission operations, interfaces, mounting & integration procedures and flow, and S/C communication, to mitigate any potential impacts to LRO’s schedule and budget.

The launch configuration has LRO sitting on top of the flight-proven B1194VS Payload Attachment Fitting (PAF), which is supported by the EELV Secondary Payload Adaptor (ESPA) ring based S-S/C. This stack sits on top of a second existing flight-proven high-capacity two-piece D1666VS PAF, which sits on the Centaur. These three elements provide a simple, direct load path. LRO’s launch umbilical is externally mounted to the S-S/C with separation connectors at the upper and lower interfaces to enable both LRO and Centaur separation from the S-S/C. Clearance is provided between LRO’s central Delta-V thrusters and support cone, which hang below the LRO-adapter separation plane and the S-S/C propellant tank and Visible and IR cameras .

Shepherd Spacecraft (S-S/C) Description
The S-S/C is an extremely simple spacecraft and consists of the components listed in Table 2.2-1, which also specifies their masses. The S-S/C has no deployments and no active mechanical elements, except for a separation band. No technology or unit developments are required except for an impact flash detection unit called a photometer. The payload consists of existing (2) visible and (4) IR cameras, (2) NIR spectrometers, and one development unit called a photometer. The entire system is low risk and benefits from LRO’s procurement and integration. The spacecraft structure is a single, ESPA ring that functions as a multifunctional integrating element which supports the LRO adapter; contains an independent set of avionics; a small 344 kg capacity monopropellant-propulsion system, a single-panel body mounted solar array and battery; and mounts for the impact observation instruments, two S-Band omni antennas, and 2 medium-gain horns.

The structural frame of the S-S/C is an ESPA ring with components mounted inside and around its exterior. The ESPA ring structure is flight qualified through analysis. Since the ESPA ring was designed to launch a primary 6800 kg spacecraft above it and up to six 200 kg spacecraft mounted radially around its cylindrical face, our usage of the ESPA ring is very low risk. LRO’s 1800 kg mass provides very high structural margins against the primary spacecraft capability of 6800 kg. The S-S/C propellant tank has a capacity of 344 kg of monopropellant hydrazine, which is still a very small load compared to the 6800 kg qualified center mount capability. The maximum load at any of the six circumferential attach points is significantly lower than 100 kg, again providing high structural margins against the ESPA ring capability. The common 62” dia. upper and lower ESPA attachments connect to the LRO-adapter/separation system and the D1666VS-PAF-interface ring/separation system, respectively. The ESPA ring itself accounts for 25% of the entire S-S/C system mass, including the attached LRO adapter. The entire S-S/C weighs ~534 kg and has over 31% dry mass margin and shown in Table 1.

The S-S/C propulsion-system tank is supported on a single cone which mounts between the upper ESPA ring mounting face and the mounting flange of the LRO adapter. The tank is an off-the-shelf Tracking & Data Relay Satellite (TDRS) tank, identical to the two being flown by LRO. The S-S/C operates with a bladder in blow-down mode, while LRO uses a separate pressurant tank to increase capacity of their TDRS tanks.

The S-S/C propulsion system uses all commercially available components. The TDRS tank is in the critical path and was ordered shortly after the Authority To Proceed (ATP) was granted to support Propulsion Integration & Test (I&T), which precedes Space Vehicle I&T. Two 5-lbf Atlantic Research Corporation thrusters are used to provide sufficient control authority over the Centaur & S-S/C stack for large delta-v maneuvers, and eight 1-lbf thrusters are used for precision attitude control at a low duty-cycle and also to provide precision pointing during the Centaur impact observation mission phase. The thrusters are arranged symmetrically around the S-S/C with 10 thrusters arranged as redundant pairs which are able to provide full 3-axis pointing and delta-V.

The propulsion distribution module and the fill & drain valves are mounted on one of the four equipment panels mounted to the radial ESPA ring attach points.

Outer radiator panels and a mounting ring are used to attach the avionics elements to the ESPA ring. Both the outer equipment mount panels and the outer radiators have simple horizontal axial groove heat pipes. C-shaped heat pipes thermally interconnect the equipment panels and the outer radiator panels. All four equipment panels have high thermal margins. Each panel’s radiator area is masked off with Multi-Layer Insulation (MLI) during I&T to match the thermal dissipation needs of the associated mounted components. The spacecraft body points the solar array to the Sun providing three high dissipation sites, two lower dissipation sites, and one Sun facing ESPA location for the solar array. The star-tracker camera and Lunar Science cameras have small dedicated radiators. The payload instrument electronics unit is mounted to one of the equipment panels.

The power system consists of the LRO Power Control Electronics (PCE), a 600W solar array, and an existing 80 A-h Li-Ion battery. Peak power for the system is 372W. The PCE can support a 1kW array and distribution. A single Output Module is all that is required for a system with a peak less than 450W. The body mounted solar array is structurally designed to be extremely high frequency and uses a large, 5”-thick honeycomb, ESPA-ring mount. The solar array is sized to provide 600W with the S-S/C and Centaur stack pointed in a +/- 10 deg Attitude Control System (ACS) dead band to the Sun. Standard 28% multi-junction cells are used in the array. With the payload off, the battery can be recharged in ~15 hours from a 50% Depth of Discharge. With the instruments on and transmitting telemetry, the battery provides nearly 2 hours of operation without charging from the solar array.

The ACS consists of a Star Tracker Assembly (STA), and Miniature Inertial Measurement Unit (MIMU), 10 Coarse Sun Sensor Assemblies (CSSA), and the Propulsion & Deployment Electronics (PDE). The ACS is based primarily on LRO hardware and software in the same single strung arrangement. LRO has selected the Honeywell MIMU gyro assembly, and this is the same unit NGST uses on STSS, but with a 1553 interface instead of our RS-422. The MIMU can be procured with built-in accelerometers to facilitate ephemeris updates by measuring delta-V increments on board. The LRO STA has not been procured yet, but it is specified to be an autonomous unit (e.g., Galileo’s A-STR), that will contain the star catalog and software to acquire and track stars, compute celestial attitude, and output a quaternion to the ACS flight software (FSW). The STA will provide 10 arc-second attitude knowledge accuracy, significantly greater than what is required for impact targeting. LRO’s CSSA suite involves 10 separate analog sensors interfacing with the Attitude Control Electronics Command & Data Handling (C&DH) Assembly. The S-S/C uses the LRO propulsion valve drive electronics repackaged in a unit called the Propulsion Deployment Electronics (PDE). The STA and MIMU provide a stellar-inertial attitude reference for all-attitude capability during all spacecraft phases. The CSSAs enable sun-pointing of the solar array initially and in the event of loss of star track by the STA. Torque and delta-V capability is provided by the Propulsion Subsystem’s N2H4 Reaction Control System (RCS) thrusters, whose valves are actuated by the PDE. Algorithms for attitude reference processing and attitude/delta-V control are implemented in the C&DH processor. All interfaces are 1553, except the CSSAs’ analog outputs, which use an A-to-D converter in the Attitude Control Electronics C&DH Assembly. The STA is mounted on the anti-sun side of the vehicle with its line of sight tilted toward the aft (Centaur) end of the vehicle. This gives us the best Earth-Sun-Moon avoidance. Reflections from the vehicle are not a problem since the STA is on its dark side.

The LCROSS communication baseline system is single-strung and can deliver 1.5 Mbps real-time data from the Moon to the Deep Space Network (DSN) 70m dish using one of the two medium gain horn antennas or can deliver 40 Kbps using one of the two omni antennas, pointed +/- 30 degrees from Earth. The S-S/C uses the existing LRO Transponder (along with other LRO passive microwave components). The Transponder provides 7W Radio Frequency (RF) output. Throughout the cruise part of the mission, the S-S/C will communicate using one of its omni antennas to a 34m dish on Earth with telemetry rates up to 40 kbps. During final lunar approach and Centaur impact observation, the spacecraft roll attitude is controlled to provide +/- 20 degrees pointing for one of the two fixed-mounted medium-gain horn antennas (~ 12 dB gain) so as to provide a 1.5 Mbps downlink when using DSN 70m dish assets. At least one of the three DSN sites has visibility to the spacecraft at all times. The White Sands Complex (WSC), which LRO plans to use as their primary ground station, has intermittent visibility (requiring comm. scheduling), approximately the same as the DSN Goldstone site. DSN 34m Stations, or WSC and other LRO backup sites around the world, could be used for routine Telemetry, Tracking, & Control (TT&C).

Avionics Subsystem and Flight Software
LCROSS flight avionics is a near build-to-print copy of the avionics developed for LRO and built at the same time, on the same Northrop Grumman production line that is currently building the LRO Engineering Development Models. Northrop Grumman is currently building the flight avionics upon which most of the LRO designs are based. No design changes are anticipated which enables the LRO qualification program to verify the mission suitability of avionics for both missions. With the exception of the new payload instruments, which interface through standard industry protocols, the avionics are simply a partially depopulated version of LRO avionics. The components removed from the LRO avionics are shown crossed-out. The C&DH subsystem is Rad-750 (radiation-hardened processor)-based and provides interfaces for 1553, RS-422, and analog signals. It also provides connectivity between the communications subsystem elements, ACS/propulsion, power and payload operation components.

To assure no impact to LRO, all testing through spacecraft I&T will be done in parallel with LRO, with no common use of test resources. A second set of Electrical Ground Support Equipment will be built and delivered to the LCROSS Spacecraft I&T team. This Leader / Follower approach has been executed before and is almost identical to the highly successful plan used for WMAP and EO-1. In the WMAP case, Northrop Grumman collaborated in the design under a Space Act Agreement and obtained the rights to Goddard’s avionics architecture. For the EO-1 project, which followed WMAP, Northrop Grumman delivered all spacecraft avionics and flight software in time to allow EO-1 to launch 7 months ahead of WMAP. LCROSS is much simpler than the EO-1 spacecraft, which supported three complex instruments requiring arc-second pointing. Four years after completing its 18 month mission, this single-string spacecraft is still performing flawlessly.

The LCROSS flight software has been incrementally developed over a series of GSFC missions, including WMAP and EO-1, and will fly on SDO and LRO. It uses a layered architecture with the hardware layer at the core, surrounded by the Operating System, which is surrounded by the communications layer with the application layer at the top of the hierarchy. The communications layer called the “Software Bus” provides for inter-processor communications, which coordinates and validates all of the inter-task message traffic between the software tasks. The application software is logically separated into several C&DH application tasks that coordinate and perform all of the Data Handling System (DHS)-required functions. Optionally, tasks required for ACS functions can be integrated into the software architecture. Each of the tasks running at the application layer performs a subset of the overall DHS requirements. Each task is responsible for implementing requirements in a particular functional area (i.e., command distribution, 1553-bus communications, and telemetry collection). The average software reuse, mission to mission is about 80%. It also makes extensive use of Tables to reconfigure from mission to mission.

Technical Information
Overview | Mission Rationale | Spacecraft and System Description | Instrumentation | Water Detection | Targeting
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